NACA 63A010 AIRFOIL (n63010a-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 63A010 AIRFOIL (n63010a-il) Reynolds number: 1,000,000 Max Cl/Cd: 64.25 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n63010a-il-1000000-n5.txt Download as CSV file: xf-n63010a-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63A010 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.000 -1.1604 0.08072 0.07868 -0.0006 1.0000 0.0041
-15.750 -1.2040 0.06763 0.06534 -0.0091 1.0000 0.0040
-15.500 -1.2307 0.05906 0.05657 -0.0145 1.0000 0.0040
-15.250 -1.2502 0.05243 0.04977 -0.0182 1.0000 0.0040
-15.000 -1.2642 0.04721 0.04437 -0.0207 1.0000 0.0041
-14.750 -1.2727 0.04313 0.04014 -0.0221 1.0000 0.0041
-14.500 -1.2773 0.03983 0.03670 -0.0228 1.0000 0.0042
-14.250 -1.2775 0.03726 0.03401 -0.0228 1.0000 0.0043
-14.000 -1.2753 0.03514 0.03178 -0.0224 1.0000 0.0044
-13.750 -1.2715 0.03331 0.02984 -0.0216 1.0000 0.0044
-13.500 -1.2661 0.03175 0.02817 -0.0205 1.0000 0.0045
-13.250 -1.2590 0.03041 0.02673 -0.0191 1.0000 0.0046
-13.000 -1.2510 0.02918 0.02540 -0.0175 1.0000 0.0047
-12.750 -1.2411 0.02806 0.02417 -0.0159 1.0000 0.0048
-12.500 -1.2282 0.02698 0.02296 -0.0146 1.0000 0.0050
-12.250 -1.2136 0.02590 0.02177 -0.0135 1.0000 0.0051
-12.000 -1.1978 0.02486 0.02061 -0.0125 1.0000 0.0053
-11.750 -1.1810 0.02385 0.01948 -0.0115 1.0000 0.0055
-11.500 -1.1631 0.02290 0.01842 -0.0106 1.0000 0.0056
-11.250 -1.1442 0.02202 0.01742 -0.0098 1.0000 0.0058
-11.000 -1.1244 0.02121 0.01651 -0.0090 1.0000 0.0059
-10.750 -1.1037 0.02048 0.01567 -0.0084 1.0000 0.0060
-10.500 -1.0850 0.01942 0.01450 -0.0074 1.0000 0.0062
-10.250 -1.0646 0.01859 0.01357 -0.0067 1.0000 0.0065
-10.000 -1.0417 0.01809 0.01304 -0.0063 1.0000 0.0068
-9.750 -1.0184 0.01765 0.01256 -0.0059 1.0000 0.0071
-9.500 -0.9950 0.01719 0.01205 -0.0054 1.0000 0.0074
-9.250 -0.9720 0.01665 0.01146 -0.0049 1.0000 0.0077
-9.000 -0.9491 0.01608 0.01082 -0.0044 1.0000 0.0079
-8.750 -0.9261 0.01553 0.01020 -0.0038 1.0000 0.0082
-8.500 -0.9029 0.01500 0.00960 -0.0032 1.0000 0.0085
-8.250 -0.8796 0.01451 0.00905 -0.0027 1.0000 0.0087
-8.000 -0.8561 0.01407 0.00855 -0.0021 1.0000 0.0089
-7.750 -0.8322 0.01371 0.00815 -0.0016 1.0000 0.0090
-7.500 -0.8116 0.01291 0.00727 -0.0005 1.0000 0.0095
-7.250 -0.7896 0.01237 0.00668 0.0004 1.0000 0.0100
-7.000 -0.7673 0.01194 0.00621 0.0013 1.0000 0.0104
-6.750 -0.7358 0.01152 0.00575 0.0002 0.9957 0.0108
-6.500 -0.7047 0.01114 0.00534 -0.0008 0.9901 0.0112
-6.250 -0.6729 0.01079 0.00497 -0.0020 0.9838 0.0118
-6.000 -0.6414 0.01046 0.00459 -0.0030 0.9753 0.0124
-5.750 -0.6104 0.01015 0.00423 -0.0039 0.9642 0.0127
-5.500 -0.5808 0.00988 0.00392 -0.0044 0.9499 0.0130
-5.250 -0.5538 0.00959 0.00356 -0.0043 0.9321 0.0137
-5.000 -0.5280 0.00935 0.00326 -0.0039 0.9130 0.0150
-4.750 -0.5024 0.00916 0.00301 -0.0034 0.8943 0.0164
-4.500 -0.4764 0.00900 0.00279 -0.0031 0.8770 0.0183
-4.250 -0.4503 0.00882 0.00257 -0.0027 0.8604 0.0235
-4.000 -0.4240 0.00865 0.00238 -0.0025 0.8451 0.0303
-3.750 -0.3975 0.00849 0.00219 -0.0023 0.8309 0.0388
-3.500 -0.3709 0.00831 0.00202 -0.0021 0.8178 0.0520
-3.000 -0.3182 0.00783 0.00168 -0.0018 0.7934 0.1077
-2.750 -0.2923 0.00749 0.00149 -0.0016 0.7824 0.1619
-2.500 -0.2670 0.00706 0.00130 -0.0014 0.7719 0.2365
-2.000 -0.2173 0.00610 0.00097 -0.0008 0.7514 0.4242
-1.750 -0.1915 0.00580 0.00088 -0.0005 0.7416 0.4937
-1.500 -0.1655 0.00557 0.00083 -0.0003 0.7316 0.5536
-1.250 -0.1383 0.00546 0.00079 -0.0002 0.7214 0.5843
-1.000 -0.1109 0.00539 0.00076 -0.0001 0.7114 0.6083
-0.750 -0.0834 0.00535 0.00074 0.0000 0.7013 0.6284
-0.500 -0.0557 0.00531 0.00073 0.0000 0.6909 0.6461
-0.250 -0.0278 0.00530 0.00072 0.0000 0.6806 0.6595
0.000 0.0000 0.00529 0.00072 0.0000 0.6704 0.6704
0.250 0.0279 0.00530 0.00072 0.0000 0.6595 0.6807
0.500 0.0557 0.00531 0.00073 0.0000 0.6461 0.6909
0.750 0.0834 0.00535 0.00074 0.0000 0.6284 0.7014
1.000 0.1109 0.00539 0.00076 0.0001 0.6082 0.7115
1.250 0.1383 0.00546 0.00079 0.0002 0.5844 0.7214
1.500 0.1655 0.00557 0.00083 0.0003 0.5537 0.7316
1.750 0.1916 0.00579 0.00088 0.0005 0.4963 0.7416
2.000 0.2174 0.00610 0.00097 0.0008 0.4255 0.7514
2.500 0.2671 0.00706 0.00130 0.0014 0.2359 0.7719
2.750 0.2924 0.00749 0.00149 0.0016 0.1610 0.7824
3.000 0.3183 0.00783 0.00168 0.0018 0.1083 0.7934
3.500 0.3710 0.00831 0.00202 0.0021 0.0519 0.8177
3.750 0.3976 0.00849 0.00219 0.0023 0.0388 0.8308
4.000 0.4241 0.00866 0.00238 0.0025 0.0303 0.8450
4.250 0.4505 0.00882 0.00257 0.0027 0.0234 0.8602
4.500 0.4765 0.00901 0.00279 0.0030 0.0183 0.8768
4.750 0.5025 0.00916 0.00301 0.0034 0.0164 0.8942
5.000 0.5282 0.00935 0.00326 0.0038 0.0149 0.9130
5.250 0.5539 0.00959 0.00356 0.0042 0.0137 0.9320
5.500 0.5809 0.00988 0.00391 0.0044 0.0130 0.9496
5.750 0.6104 0.01015 0.00423 0.0039 0.0127 0.9639
6.000 0.6414 0.01046 0.00459 0.0030 0.0123 0.9752
6.250 0.6730 0.01079 0.00497 0.0019 0.0118 0.9837
6.500 0.7048 0.01114 0.00534 0.0008 0.0112 0.9902
6.750 0.7359 0.01152 0.00575 -0.0002 0.0108 0.9958
7.000 0.7671 0.01194 0.00621 -0.0013 0.0103 1.0000
7.250 0.7895 0.01237 0.00668 -0.0004 0.0100 1.0000
7.500 0.8116 0.01291 0.00727 0.0005 0.0095 1.0000
7.750 0.8321 0.01372 0.00816 0.0016 0.0090 1.0000
8.000 0.8562 0.01407 0.00855 0.0021 0.0089 1.0000
8.250 0.8797 0.01452 0.00905 0.0026 0.0087 1.0000
8.500 0.9032 0.01501 0.00960 0.0032 0.0085 1.0000
8.750 0.9264 0.01553 0.01021 0.0038 0.0082 1.0000
9.000 0.9494 0.01609 0.01083 0.0043 0.0080 1.0000
9.250 0.9724 0.01666 0.01147 0.0049 0.0077 1.0000
9.500 0.9955 0.01719 0.01206 0.0054 0.0074 1.0000
9.750 1.0188 0.01766 0.01257 0.0058 0.0071 1.0000
10.000 1.0423 0.01808 0.01303 0.0062 0.0068 1.0000
10.250 1.0651 0.01860 0.01358 0.0066 0.0065 1.0000
10.500 1.0855 0.01943 0.01450 0.0073 0.0062 1.0000
10.750 1.1043 0.02049 0.01568 0.0082 0.0060 1.0000
11.000 1.1250 0.02122 0.01652 0.0089 0.0059 1.0000
11.250 1.1449 0.02202 0.01743 0.0096 0.0058 1.0000
11.500 1.1638 0.02291 0.01842 0.0104 0.0056 1.0000
11.750 1.1817 0.02386 0.01949 0.0113 0.0055 1.0000
12.000 1.1986 0.02487 0.02062 0.0123 0.0053 1.0000
12.250 1.2144 0.02591 0.02177 0.0133 0.0051 1.0000
12.500 1.2291 0.02698 0.02296 0.0144 0.0050 1.0000
12.750 1.2421 0.02807 0.02418 0.0157 0.0048 1.0000
13.000 1.2520 0.02919 0.02541 0.0173 0.0047 1.0000
13.250 1.2602 0.03041 0.02673 0.0189 0.0046 1.0000
13.500 1.2675 0.03173 0.02815 0.0203 0.0045 1.0000
13.750 1.2729 0.03329 0.02982 0.0214 0.0044 1.0000
14.000 1.2770 0.03509 0.03173 0.0222 0.0044 1.0000
14.250 1.2790 0.03724 0.03399 0.0226 0.0043 1.0000
14.500 1.2786 0.03983 0.03670 0.0225 0.0042 1.0000
14.750 1.2746 0.04306 0.04007 0.0219 0.0041 1.0000
15.000 1.2658 0.04718 0.04435 0.0205 0.0041 1.0000
15.250 1.2518 0.05242 0.04976 0.0180 0.0040 1.0000
15.500 1.2325 0.05905 0.05657 0.0142 0.0040 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 63A010 AIRFOIL (n63010a-il)